Aerial vehicle

ABSTRACT

Aircraft capable of vertical takeoff and landing, hovering, and efficient forward flight are described. An aircraft includes two side mounted tiltable proprotors and a central rotor disposed above the proprotors. The proprotors are tiltable between at least a horizontal position for forward flight and a vertical position for vertical or hovering flight. The central rotor may be powered for vertical and transitional flight modes and may turn by free autorotation during forward flight. The proprotors may be differentially tilted during vertical or hovering flight to counter torque effects of the central rotor. The central rotor may be foldable and/or easily detachable from the aircraft to facilitate storage and transportation. Left and right proprotors may provide both forward thrust and attitude control. Control inputs to left and right proprotors may be connected directly to an autopilot creating closed loop actuation using motor RPM feedback.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. application Ser. No.17/662,217, filed May 5, 2022, entitled “AERIAL VEHICLE,” which is acontinuation of U.S. application Ser. No. 17/461,719, filed Aug. 30,2021, entitled “AERIAL VEHICLE,” which is a continuation-in-part ofInternational PCT Application No. PCT/US2020/046240, filed Aug. 13,2020, entitled “AERIAL VEHICLE,” which claims the benefit of U.S.Provisional Application Ser. No. 62/886,578, filed Aug. 14, 2019,entitled “AERIAL VEHICLE,” U.S. Provisional Application Ser. No.62/896,257, filed Sep. 5, 2019, entitled “AERIAL VEHICLE,” and U.S.Provisional Application Ser. No. 63/018,848, filed May 1, 2020, entitled“AERIAL VEHICLE,” all of which are incorporated by reference in theirentirety and for all purposes.

FIELD

The present disclosure relates generally to manned or unmanned aircraft,and more particularly to aircraft capable of vertical and horizontalflight.

BACKGROUND

A helicopter is an aircraft in which lift and thrust are supplied by oneor more horizontal rotors. The advantages of a helicopter include itsability to hover and to take off and land vertically. However, amongother things, a helicopter suffers from its relatively poor operatingenergy efficiency compared to fixed-wing aircraft.

A gyroplane (also known as a gyrocopter or autogyro) is an aircraft thatuses an unpowered rotor in autorotation to develop lift. Autorotation isa rotor state in which the rotor derives from the freestream 100% of thepower required to rotate it, and the resulting rotation provides lift.In a gyrocopter, forward thrust is typically provided by anengine-driven propeller. However, like a fixed-wing aircraft, agyrocopter cannot take off and land vertically.

In the case of a gyroplane, as the aircraft goes down the runway andgathers speed, the overhead rotor's shaft is tilted backwards allowingthe wind to blow though the rotor in order to start turning. As therotor reaches a certain RPM, that very rotor becomes a “virtual wing”for the gyrocopter providing lift. Once it reaches the desired RPM, thegyrocopter is ready for takeoff. As the gyrocopter gains speed in theair the angle of attack of the rotor head (virtual wing) is reduced bymoving the shaft forward, increasing speed and reducing drag. Agyrocopter does not have the ability to achieve vertical takeoff becauseit does not have a variable pitch rotor. The pitch of the blades of agyrocopter is typically zero. In addition to being able to tilt theshaft forwards and backwards, a gyrocopter can also tilt the shaft/rotorhead to starboard and port, providing aileron like maneuverability.Also, for a gyrocopter to fly in a stable manner, it has a unique rotorhead assembly that can teeter totter on the shaft allowing the bladesfreedom of movement as they rotate.

A helicopter has a fixed vertical shaft with swash plates and links tothe rotor head, allowing the pilot to modify the pitch of the rotorblades, generating lift. Further, it has a propeller mounted verticallyat the end of a boom (or a ducted fan), creating a thrust to counter thetorque generated by the motor that drives the rotor head.

In the case of a helicopter the blades drive the air from the topdownwards creating thrust and lift. In the case of a gyrocopter the airflows through the blades upwards spinning them and creating lift.

SUMMARY

The systems, methods, and devices of the present technology each haveseveral innovative aspects, no single one of which is solely responsiblefor its desirable attributes disclosed herein. Without limiting thescope of this disclosure, its more prominent features will now bediscussed briefly.

In a first aspect, an aircraft comprises a fuselage, a mast extendingupward from the fuselage and fixed to the fuselage at a lower end of themast, a rotor rotatably coupled to an upper end of the mast opposite thelower end, a rotor motor disposed at the upper end of the mast andconfigured to cause rotation of the rotor in a first direction, alateral boom coupled to an intermediate section of the mast between theupper end and the lower end, a first powered proprotor disposed at afirst end of the lateral boom, a second powered proprotor disposed at asecond end of the lateral boom opposite the first end, a first proprotortilt servo configured to control an orientation the first poweredproprotor between at least a horizontal tilt angle and a vertical tiltangle, and a second proprotor tilt servo configured to control anorientation of the second powered proprotor between at least thehorizontal tilt angle and the vertical tilt angle independent of theorientation of the first powered proprotor.

In some embodiments, the first powered proprotor is powered by a firstproprotor motor, the second powered proprotor is powered by a secondproprotor motor, and rotational speeds of the first proprotor motor andthe second proprotor motor are independently variable. In someembodiments, the first and second proprotor motors are configured tocounteract torque effects of the rotor motor during vertical or hoveringflight by powering the first and second proprotors at different speeds.In some embodiments, an output of the rotor motor is rotationallycoupled to the rotor by a one-way bearing that transmits torque from therotor motor to the rotor when the rotor motor is activated and allowsthe rotor to turn freely in the first direction when the rotor motor isdeactivated. In some embodiments, the first and second proprotor tiltservos are configured to control the orientations of the first andsecond powered proprotors within a range of tilt angles of greater than90 degrees. In some embodiments, the range of tilt angles isapproximately 150 degrees. In some embodiments, the first and secondpowered proprotors are tiltable between a first extreme tilt angle of atleast 25 degrees aft of vertical and a second extreme tilt angle of atleast 25 degrees below horizontal. In some embodiments, the firstextreme tilt angle is at least 30 degrees aft of vertical, and thesecond extreme tilt angle is at least 30 degrees below horizontal. Insome embodiments, the aircraft further comprises a rotor tilt servoconfigured to tilt the rotor between at least a first tilt angle inwhich the rotor spins about a vertical axis of the aircraft and a secondtilt angle in which the rotor spins about an axis angled atapproximately 20 degrees relative to the vertical axis of the aircraft.In some embodiments, the first and second proprotor tilt servos areconfigured to counteract torque effects of the rotor motor duringvertical or hovering flight by differentially tilting the first andsecond powered proprotors relative to a vertical axis of the aircraft.In some embodiments, the aircraft is configured to fly in a plurality offlight configurations including a vertical flight configuration in whichthe first and second powered proprotors are disposed within 30 degreesof vertical and the rotor is driven by the rotor motor, and a horizontalflight configuration in which the first and second powered proprotorsare disposed within 30 degrees of horizontal and the rotor turns by freeautorotation. In some embodiments, the rotor comprises rotor bladesconfigured to automatically adjust to a positive blade pitch in thevertical flight configuration and to a flat blade pitch in thehorizontal flight configuration. In some embodiments, the aircraftfurther comprises a tiltable empennage including a horizontalstabilizer, the tiltable empennage configured to rotate about alengthwise axis of the horizontal stabilizer. In some embodiments, thetiltable empennage is configured to rotate to a lowered position duringvertical flight such that the horizontal stabilizer is aligned with adownwash created by the rotor when the rotor motor is engaged.

In a second aspect, an aircraft comprises a rotor comprising a pluralityof rotor blades, a lower rotor hub, and an upper rotor hub assemblyincluding at least one mounting member. The lower rotor hub comprises arotor mount shaft extending along an axis of rotation of the rotor, andat least one mounting pin hole extending through the rotor mount shaft.Each mounting member comprises a central section having two mounting pinholes extending therethrough such that the central section can becoupled to the lower rotor hub by inserting a mounting pin through thetwo mounting pin holes of the central section, and two mounting bracketsdisposed at opposite ends of the central section, each mounting bracketfixedly coupled to one of the plurality of rotor blades.

In some embodiments, the upper rotor hub assembly further comprises ablade pitch adjustment linkage configured to automatically adjust ablade pitch of the rotor blades based on a rotational velocity of therotor. In some embodiments, the blade pitch adjustment linkage comprisesat least one biasing element configured to cause an increase in theblade pitch when subjected to an increased centrifugal force associatedwith an increase in the rotational velocity of the rotor. In someembodiments, the biasing element comprises a gas spring coupled to thecentral section and one of the two mounting brackets. In someembodiments, each mounting bracket of each mounting member is coupled tothe central section by a hinge allowing the mounting member to be foldedbetween an extended configuration and a folded configuration, eachmounting bracket further comprising a plurality of locking pin holesconfigured to lock the mounting member in the extended configurationwhen a locking pin is disposed within the locking pin holes. In someembodiments, each mounting member can teeter about the mounting pin. Insome embodiments, the aircraft comprises at least four blades and atleast two mounting members, each mounting member being coupled to twooppositely disposed rotor blades of the plurality of rotor blades, andeach mounting member is shaped to permit at least one other mountingmember to teeter independently. In some embodiments, each rotor blade isslidably attachable and detachable from the mounting brackets. In someembodiments, each mounting bracket of each mounting member is attachedto a mounting body sized and shaped to fit within a mounting bodyopening of the mounting member, and the mounting body is securable tothe mounting member by inserting a blade attachment pin through pinholes in the mounting member and the mounting body. In some embodiments,the upper rotor hub assembly further comprises an upper rotor mount hubcomprising a tubular shaft at least partially surrounding the rotormount shaft of the lower rotor mount hub, wherein each mounting memberis coupled to the upper rotor mount hub by at least one mounting pin,and wherein the tubular shaft is detachably coupled to the lower rotormount shaft by at least one mounting pin.

In a third aspect, a rotor assembly for an aircraft comprises aplurality of rotor blades and an upper rotor hub assembly including atleast one mounting member. Each mounting member comprises a centralsection; two mounting brackets disposed at opposite ends of the centralsection, each mounting bracket fixedly coupled to one of the pluralityof rotor blades; and a blade pitch adjustment linkage coupled to themounting brackets and the central section, the blade pitch adjustmentlinkage configured to automatically adjust a blade pitch of the rotorblades based on a rotational velocity of the rotor blades.

In some embodiments, the blade pitch adjustment linkage comprises atleast one biasing element configured to cause an increase in the bladepitch when subjected to an increased centrifugal force associated withan increase in the rotational velocity of the rotor blades. In someembodiments, the biasing element comprises a gas spring coupled to thecentral section and one of the two mounting brackets. In someembodiments, the blade pitch adjustment linkage includes at least onesynchronization linkage coupled to the mounting brackets and configuredto synchronize the blade pitch of the rotor blades.

In a fourth aspect, a rotor assembly for an aircraft comprises a firstand second rotor blade, and a mounting member to which the first andsecond rotor blades are coupled in an opposed arrangement. The mountingmember comprises a central section comprising a pivot coupler configuredto allow the mounting member and rotor blades to teeter relative to theaircraft; two mounting brackets disposed at opposite ends of the centralsection, each mounting bracket fixedly coupled to first and second rotorblades; and a blade pitch linkage coupled to the mounting brackets andthe central section, the blade pitch linkage configured to synchronize apitch angle of the first and second rotor blades.

In some embodiments, the rotor assembly comprises first and second innercylinders extending from the central section, wherein the mountingbrackets comprise first and second outer cylinders that slidinglyreceive the first and second inner cylinders, wherein the outercylinders comprise pin-receiving apertures and the inner cylinderscomprise apertures defining a track, wherein the blade pitch linkagecomprises pins extending through the apertures in the first and secondcylinders such that when the outer cylinders slide relative to the innercylinders, the pins slide through the tracks, thereby causing rotationof the outer cylinders relative to the inner cylinders.

BRIEF DESCRIPTION OF THE DRAWINGS

The above-mentioned aspects, as well as other features, aspects, andadvantages of the present technology will now be described in connectionwith various implementations, with reference to the accompanyingdrawings. The illustrated implementations are merely examples and arenot intended to be limiting. Throughout the drawings, similar symbolstypically identify similar components, unless context dictatesotherwise.

FIG. 1 depicts an isometric view of an aircraft in accordance with anexample embodiment of the present technology.

FIG. 2 depicts a side view of the aircraft of FIG. 1 .

FIG. 3 depicts a rear isometric view of the aircraft of FIGS. 1 and 2 .

FIG. 4 depicts a top front view of the aircraft of FIGS. 1-3 .

FIG. 5 schematically illustrates an example tilting range of proprotorswith respect to the lateral boom of an example aircraft.

FIG. 6 schematically illustrates example VTOL/hover and forward flighttilt positions of the proprotors.

FIG. 7 depicts a medial portion of the lateral boom of an exampleaircraft including proprotor support and control components locatedthereon.

FIG. 8 is a detailed view of certain proprotor support and controlcomponents of FIG. 7 .

FIG. 9 depicts medial and lateral portions of the lateral boom of theexample aircraft of FIGS. 7 and 8 including proprotor support andcontrol components located thereon.

FIG. 10 depicts a lateral portion of the lateral boom of the exampleaircraft of FIGS. 7-9 including proprotor support and control componentslocated thereon.

FIG. 11 depicts a side view of a lower rotor hub of an example aircraft.

FIG. 12 depicts a rear view of the lower rotor hub of the exampleaircraft of FIG. 11 .

FIG. 13 depicts a top perspective view of the lower rotor hub of theexample aircraft of FIGS. 11 and 12 .

FIG. 14 depicts a side view of the lower rotor hub of the exampleaircraft of FIGS. 11-13 in a vertical configuration.

FIG. 15 depicts a side view of the lower rotor hub of the exampleaircraft of FIGS. 11-14 in a tilted configuration.

FIG. 16 depicts an upper rotor hub assembly including two mountingmembers coupleable to the lower rotor hub of FIGS. 11-15 .

FIGS. 17 and 18 depict a rotor blade mounted to a blade arm of amounting member in accordance with an example embodiment.

FIGS. 19 and 20 depict the upper rotor hub assembly in a foldedconfiguration.

FIG. 21 depicts a partial isometric view of the upper rotor hub assemblyof FIGS. 16-20 coupled to the lower rotor hub of FIGS. 11-15 .

FIG. 22 depicts a partial top view of the upper rotor hub assembly ofFIGS. 16-20 coupled to the lower rotor hub of FIGS. 11-15 .

FIG. 23 depicts a perspective view of a portion of an upper rotor hubassembly in accordance with some embodiments.

FIG. 24 depicts a side view of the portion of an upper rotor hubassembly of FIG. 23 coupled to the lower rotor hub of an exampleaircraft.

FIGS. 25 and 26 depict an example embodiment of an upper rotor hubassembly with a bayonet-type blade mounting system.

FIG. 27 depicts an example embodiment of a rotor hub assembly includinga teetering pivot point.

FIGS. 28 and 29 illustrate side and cross-sectional views of the rotorhub assembly of FIG. 27 in a zero pitch configuration.

FIGS. 30 and 31 illustrate side and cross-sectional views of the rotorhub assembly of FIGS. 27-29 in a positive pitch configuration.

FIG. 32 depicts an exploded view of the rotor hub assembly of FIGS.27-31 .

FIG. 33 illustrates an example gearing system for turning the upperrotor hub assembly of FIGS. 27-32 .

FIG. 34 is a cross-sectional view illustrating a braking system forreducing the rotational speed of the upper rotor hub assembly of FIGS.27-32 .

FIGS. 35-37 illustrate an example aircraft including a tiltableempennage.

FIGS. 38-41 illustrate an example aircraft including a single proprotorand pivotable wings configured to operate as control surfaces.

FIGS. 42 and 43 illustrate a further example configuration of anaircraft in accordance with the present technology.

FIGS. 44 and 45 illustrate a further example configuration of a rotorhead assembly in accordance with the present technology.

DETAILED DESCRIPTION

Generally described, embodiments of the present disclosure provideaircraft as well as aircraft systems, components, and control methodsproviding enhanced flight characteristics relative to existinghelicopters, gyroplanes, fixed-wing airplanes, and tiltrotor aircraft.Aircraft disclosed herein may be capable of efficient forward flight,hovering, and/or vertical takeoff and landing (VTOL), as well astransitioning between flight modes in flight. For example, an aircraftin accordance with the present technology may be able to take offvertically from a deployment location, transition to a forward flightmode for generally horizontal flight to a remote location, transition toa vertical flight mode to hover at the remote location for an extendedtime period, transition to the forward flight mode for generallyhorizontal flight to a landing location (e.g., the deployment location),and transition again to the vertical flight mode to land at the landinglocation. Thus, the present disclosure provides aerial vehicles capableof extended data gathering or observation at a relatively distantlocation beyond the range of a traditional helicopter, without requiringa runway for takeoff and landing.

Some aircraft described herein are further configured to be partiallydisassembled and may include foldable components to provide a morecompact configuration for transportation to or from deploymentlocations. In one example, a central rotor of the aircraft may includeone or more pairs of blades attached to an upper rotor hub assembly. Theupper rotor hub assembly may be detachable from a lower rotor hub of theaircraft, and may permit the rotor blades to be folded into asubstantially parallel configuration such that the central rotor may betransported in a container approximately the same length as anindividual rotor blade, without requiring the rotor blades to beindividually separated from the upper rotor hub assembly. When theaircraft is to be deployed again, the central rotor may be convenientlyunfolded and attached to the lower rotor hub without requiringadditional calibration, alignment, and the like.

FIGS. 1-4 illustrate an example aircraft 100 configured for in-flighttransition between vertical and horizontal flight modes. The aircraft100 includes a fuselage 102, a mast 120 extending generally upward fromthe fuselage 102, and an empennage 150 disposed aft of the fuselage 102.A tiltable central rotor 140 is rotatably coupled at an upper end of themast 120. Tiltably mounted proprotors 136 l, 136 r are rotatably coupledat opposing ends of a boom 132 extending laterally from the mast 120.

The fuselage 102 is a body section of the aircraft 100 and may includean interior volume sized and shaped to hold a payload. For example, theinterior volume of the fuselage 102 may be used to contain one or moreitems being transported by the aircraft 100. In some embodiments, thefuselage 102 may contain one or more reconnaissance or surveillancedevices, such as imaging devices (e.g., a visible light camera, infraredcamera, thermal camera, still camera, video camera, synthetic-apertureradar, etc.), listening devices, communications devices, or the like.The fuselage 102 may further contain at least some of the controlsystems for the aircraft 100, such as motor, control surface, or tiltservo controllers, autopilot systems, and the like. For example, in someembodiments, an autopilot system may be connected to proprotor speed andtilt control circuitry such that the autopilot can directly control thespeed, direction of rotation, and/or tilt of the proprotors 136 l, 136 ras pitch, roll, and/or yaw control devices. If the aircraft 100 isconfigured for operation as a remotely piloted unmanned aerial vehicle(UAV) or drone, the fuselage 102 may also include a communication systemto receive control commands from a remote pilot. An undercarriage 104may be disposed on a side or bottom portion of the fuselage 102 for useduring takeoff and landing phases of flight, and may include wheeledlanding gear, skids, and/or any other suitable type of undercarriage.The fuselage 102 and the undercarriage 104 may comprise any suitablyrigid or semi-rigid materials such as metal, plastic, carbon fiber,wood, fiberglass, etc.

The mast 120 extends generally upward from the fuselage 102 and supportsthe central rotor 140 disposed at an upper end of the mast 120. In someembodiments, such as in the embodiment illustrated in FIGS. 1-4 , themast also includes a forward tilt such that the mast extends upward at aforward angle from the fuselage 102. As shown in FIG. 2 , the forwardtilt of the mast 120 may advantageously allow the rotational axis of thecentral rotor 140 to align with the center of gravity of the fuselage102. In addition, when the proprotors 136 l, 136 r are tilted upward forvertical/hover flight, the central rotor 140, the longitudinal locationof the proprotors 136 l, 136 r, and the center of gravity of thefuselage 102 are all aligned with a common center of gravity forimproved stability in hover or vertical flight. The mast 120 may alsoserve as a central attachment point for other components of the aircraft100, including the boom 132, empennage 150, and lower rotor hub 300. Themast 120 may further include an interior volume in which one or moreaircraft components may be disposed. In some embodiments, energy storagemedia such as batteries, hydrogen storage, or the like, may be containedwithin the mast 120. Placement of batteries or hydrogen storage withinthe mast 120 may advantageously keep such relatively heavy componentsclose to the center of gravity of the aircraft 100, resulting inimproved stability.

The central rotor 140 is rotatably mounted to a top portion of the shaft120 via a lower rotor hub 300. Rotor blades 142 are fixed to an upperrotor hub assembly 400 coupled to the lower rotor hub 300. The rotorblades 142 may have an airfoil profile configured to enhance theproduction of lift while the central rotor 140 is spinning. Although therotor 140 of the example aircraft of FIGS. 1-4 has four rotor blades142, it will be understood that other numbers of rotor blades 142 may beincluded. For example, in some embodiments, the rotor 140 may have 2, 3,4, 5, 6, or more rotor blades 142.

The lower rotor hub 300 is tiltably mounted at the top of the mast 120,for example, on a central rotor control housing 122. The central rotorcontrol housing 122 may include one or more servos configured to provideat least fore and aft tilting of the central rotor 140 (e.g., by tiltingthe lower rotor hub 300). In some embodiments, the central rotor controlhousing 122 further includes one or more servos configured to providelateral tilting of the central rotor 140. The central rotor controlhousing 122 can also include a rotor motor configured to turn thecentral rotor 140 during some phases of flight (e.g., during verticaland/or transitional flight modes), as described in greater detail below.In some embodiments, the rotor motor may be configured to turn thecentral rotor 140 at rotational speeds up to typical gyroplane rotorrotational speeds (e.g., up to approximately 200-600 rpm), which may beslower than typical helicopter rotor rotational speeds (e.g.,approximately 400-1500 rpm or more). The rotor motor may be rotationallycoupled to the central rotor 140 by a clutch and/or a one-way bearingsuch that, during forward flight, the central rotor 140 can rotatefaster than the rotor motor, and such that the central rotor 140 cancontinue rotating by autorotation when the rotor motor is not turning.

The proprotors 136 l, 136 r are configured to provide lift and/orforward thrust, depending on the tilt of the proprotors 136 l, 136 r.The proprotors 136 l, 136 r are tiltably mounted at opposite ends of theboom 132. In some embodiments, the boom 132 includes distal arms 134 land 134 r, which are individually pivotable along the lateral axis ofthe boom, such that each proprotor 136 l, 136 r is independentlytiltable by pivoting the distal arms 1341, 134 r of the boom 132. Servosand/or other actuators disposed within a proprotor tilt control housing130 may control pivoting of the distal arms 134 l, 134 r. As describedin greater detail below, each proprotor 136 l, 136 r can beindependently powered by left and right proprotor motors and may beoperable at different relative speeds for enhanced maneuverability. Forexample, as described above, the speed, direction of rotation, and/ortilt of each proprotor 136 l, 136 r may be independently controlled suchas by an autopilot or other control system such that the proprotors 136l, 136 r can function as control devices to control roll, pitch, and/oryaw of the aircraft 100 without requiring conventional control surfaces.In another example, attitude of the aircraft may be controlled bycontrolling the thrust of the proprotors 136 l, 136 r together incoordination with tilt of the proprotors 136 l, 136 r. The speeds ofboth proprotors 136 l, 136 r may further be adjusted collectively so asto propel the aircraft forward at a range of desired speeds. In someembodiments, the proprotors 136 l, 136 r may have blades featuring ahybrid shape between a propeller shape and a rotor shape so as tooperate efficiently in both forward and vertical/hover flight modes.

The empennage 150 is mounted at a rear portion of the aircraft 100 andincludes a horizontal stabilizer 152 and vertical stabilizers 154. Alongitudinal tail boom 156 may fix the empennage 150 to the mast 120 orfuselage 102. The horizontal stabilizer 152 and vertical stabilizers 154provide stability to the aircraft, primarily during horizontal flight.In some embodiments, the empennage 150 may include one or more controlsurfaces such as an elevator disposed on the horizontal stabilizer 152and/or rudders disposed on the vertical stabilizers 154. In embodimentshaving rudders on the vertical stabilizers 154, the vertical stabilizers154 may be placed at a location within the slipstreams of the proprotors136 l, 136 r to increase rudder effectiveness at low airspeeds. However,it will be understood that in various embodiments, the need for controlsurfaces may be eliminated by the use of independently controllable andtiltable proprotors 136 l, 136 r as described above.

With reference to FIGS. 5 and 6 , and with continued reference to FIGS.1-4 , example flight control systems and methods will now be described.FIG. 5 schematically illustrates an example tilting range of proprotors136 l, 136 r with respect to the lateral boom 132 of the aircraft 100.FIG. 6 schematically illustrates example VTOL/hover and forward flighttilt positions of the proprotors. Each of the proprotors 136 l, 136 rmay have a range of tilt angles of at least 90 degrees, and in somecases up to 150 degrees or more. For example, proprotors 136 l, 136 rmay be tiltable to a VTOL/hover position or range of positions, in whicheach proprotor 136 l, 136 r is substantially aligned about a vertical orz-axis to produce an upward lifting force. Proprotors 136 l, 136 r mayfurther be tiltable to a forward flight or horizontal flight position orrange of positions, in which each proprotor 136 l, 136 r issubstantially aligned about a longitudinal or x-axis to produce forwardthrust. Within either the VTOL/hover position or forward flightposition, each proprotor 136 l, 136 r may be independently tiltablewithin a range of approximately 15 degrees, 20 degrees, 25 degrees, 30degrees, or more, such as for maneuvering and/or stability as describedbelow. In some embodiments, the proprotors 136 l, 136 r can be tilted toany tilt angle between 30 degrees below horizontal to 30 degrees aft ofvertical, for a full tilt angle range of approximately 150 degrees.

Vertical flight modes, such as VTOL, hovering, may be achieved with theproprotors 136 l, 136 r in a VTOL/hover position, such as wherein theaxis of rotation of each of the proprotors 136 l, 136 r is withinapproximately 30 degrees of vertical. In the VTOL/hover position, thespinning proprotors 136 l, 136 r primarily generate an upward liftingforce. In addition, the central rotor 140 may also be used duringvertical flight modes. The central rotor 140 may be maintained in asubstantially vertical orientation and turned by the rotor motor at aspeed sufficient to provide gyroscopic stabilization to the aircraft 100and produce an additional lifting force in addition to the lift providedby the proprotors 136 l, 136 r. For example, turning the central rotor140 at a relatively low rotational speed (e.g., a typical gyroplanerotor rotational speed) can be sufficient to significantly stabilize theaircraft 100. However, in addition to stabilizing the aircraft 100, thecentral rotor 140 when powered may also create torque effects andgyroscopic precession. In some embodiments, the proprotors 136 l, 136 rmay be differentially controlled, and/or may be configured to rotate ina direction opposite the rotation of the central rotor 140, tocounteract the torque effect of the central rotor 140. For example, thetorque generated by a clockwise turning central rotor 140 may tend tocause the body of the aircraft 100 to spin counterclockwise. Tocounteract this torque effect, the left proprotor 136 l may be tiltedslightly forward of vertical while the right proprotor 136 r is tiltedslightly aft of vertical, such that the proprotors 136 l, 136 r exert aclockwise torque on the body of the aircraft 100 (e.g., a yawing moment)that counteracts the torque effect of the turning central rotor 140. Insome embodiments, a yawing moment may be created by varying the speedsand/or rotational directions of the proprotors 136 l, 136 r. Forexample, rotating proprotor 136 l faster than proprotor 136 r creates ayawing moment to the right, and rotating proprotor 136 r faster thanproprotor 136 l creates a yawing moment to the left. These methods ofgenerating a yawing moment may be utilized for turning and/or forcountering the torque generated when the main rotor is powered.

During vertical flight, such as takeoff, landing, hovering, lateralmovement, and/or slow forward or backward flight, differential controlof the proprotors 136 l, 136 r may further be used to maneuver theaircraft 100. As described above, differential tilting, motor speed, orrotational direction of the proprotors 136 l, 136 r may be used toproduce a yawing moment for maneuverability about the vertical axis.Differential rotation speed, rotational direction, and/or tilt of theproprotors 136 l, 136 r may be used to produce a rolling moment formaneuverability about the longitudinal axis. For example, powering theleft proprotor 136 l at a higher rotational speed relative to the rightproprotor 136 r produces a right or clockwise rolling moment; poweringthe right proprotor 136 r at a higher rotational speed relative to theleft proprotor 136 l produces a left or counterclockwise rolling moment.Simultaneous tilting of the proprotors 136 l, 136 r forward or aft ofvertical, and/or tilting the central rotor 140 forward or aft, producesa pitching moment for maneuverability about the lateral axis. Rolling orpitching the aircraft 100 out of a vertical orientation yields ahorizontal component of lift which may be utilized for forward,backward, and/or lateral movement during generally vertical flight.

Horizontal or forward flight, including straight and level flight,climbing, descending, turning, and the like, may be achieved with theproprotors 136 l, 136 r in a forward flight position, such as whereinthe axis of rotation of each proprotor 136 l, 136 r is withinapproximately 30 degrees of horizontal. In the forward flight position,the spinning proprotors 136 l, 136 r primarily generate forward thrustsubstantially parallel to a direction of flight. During forward flight,the central rotor 140 may be unpowered and turns in free autorotation.Preferably, the central rotor 140 has an upward tilt (e.g. betweenapproximately 1 degree and 20 degrees or more with the higher side ofthe central rotor 140 oriented toward the direction of flight). In someembodiments, the central rotor 140 is tilted at an angle ofapproximately 1 to 20 degrees in forward flight. Thus, in forwardflight, the aircraft 100 performs substantially as a gyroplane, with thepowered proprotors 136 l, 136 r generating thrust and the autorotatingcentral rotor 140 providing lift. The empennage 150 provides directionalstability during forward flight.

In various embodiments, the empennage 150 may or may not include controlsurfaces. For example, the horizontal stabilizer 152 may include one ormore elevators configured to provide pitch control, and the verticalstabilizers 154 may each include a rudder configured to provide yawcontrol. In some embodiments, the empennage 150 includes only rudders oronly an elevator, and in other embodiments the empennage 150 containsneither rudders nor elevators. Instead, any or all of pitch, yaw, androll may be controlled by the variable tilt and pitch proprotors. Pitchand/or roll may also be controlled at least in part by tilting of thecentral rotor 140.

Pitch control in forward flight may be achieved by tilting the centralrotor 140 forward or aft. Pitch control in forward flight may also beachieved by simultaneously tilting both proprotors 136 l, 136 r higheror lower relative to the longitudinal or x-axis. For example,simultaneous upward tilting of the proprotors 136 l, 136 r can produce anose-up pitching moment, and simultaneous downward tilting of theproprotors 136 l, 136 r can produce a nose-down forward pitching moment.

Roll control in forward flight may be achieved by tilting the centralrotor 140 left or right. However, controlling roll by tilting thecentral rotor 140 requires a lateral tilting mechanism for the centralrotor 140. In some embodiments, the lower rotor hub 300 and/or the upperrotor hub assembly 400 may be simplified by providing only fore and afttilting, and not lateral tilting. In such embodiments, aircraft roll canbe achieved based on differential tilting of the proprotors 136 l, 136r. For example, tilting the left proprotor 136 l slightly upwardrelative to horizontal and/or tilting the right proprotor 136 r slightlydownward relative to horizontal produces a right or clockwise rollingmoment. Tilting the right proprotor 136 r slightly upward relative tohorizontal and/or tilting the left proprotor 136 l slightly downwardrelative to horizontal produces a left or counterclockwise rollingmoment.

Yaw control in forward flight may be achieved by providing differentialpower to the left and right proprotors 136 l, 136 r. For example,varying the relative speeds of the proprotors 136 l, 136 r such that theleft proprotor 136 l turns at a higher rotational velocity than theright proprotor 136 r produces a right yawing moment. Varying therelative speeds of the proprotors 136 l, 136 r such that the rightproprotor 136 r turns at a higher rotational velocity than the leftproprotor 136 l produces a left yawing moment.

In various embodiments, turning in forward flight may be achieved by yawcontrol (e.g., by variable relative proprotor speeds), by roll controlonly (e.g., by variable relative tilt of the proprotors), or by acombination of yaw control and roll control. In some embodiments,turning with a combination of yaw and roll control may be desirable inorder to maintain coordinated flight without sideslip while turning. Forexample, a right turn may be performed by simultaneously (orsubstantially simultaneously) tilting the left proprotor 136 l upward,increasing the rotational speed of the left proprotor 136 l, tilting theright proprotor 136 r downward, and/or decreasing the rotational speedof the right proprotor 136 r.

In addition to vertical and forward flight modes, the aircraft 100 isfurther capable of transitioning between forward and vertical flightmodes while in flight. Maneuvers for transitioning from vertical flightto forward flight, and from forward flight to vertical flight, will nowbe described. Advantageously, the aircraft of the present disclosure arecapable of transitioning seamlessly between flight modes withoutsacrificing stability or controllability during the transition.

The aircraft 100 may transition from vertical flight to forward flightat various times during a mission, for example, after a vertical takeoffwhen entering a cruise portion of flight to a remote location, after aperiod of hovering at the remote location, etc. The transition fromvertical flight to forward flight begins with the aircraft 100configured for vertical flight. In this configuration, the proprotors136 l, 136 r are in the VTOL/hover position illustrated in FIG. 6 , andthe central rotor 140 may be turning under power while substantiallyparallel to the proprotors 136 l, 136 r. To transition to forwardflight, the proprotors 136 l, 136 r simultaneously tilt forward into theforward flight position illustrated in FIG. 6 . At approximately thesame time, the central rotor 140 is tilted rearward (e.g., higher towardthe front of the aircraft and lower toward the rear of the aircraft). Ifthe central rotor 140 was powered during the preceding vertical orhovering flight, the rotor motor may continue to turn the central rotor140 until the aircraft is established in forward flight. If the centralrotor 140 was not powered during the preceding vertical or hoveringflight, the rotor motor may activate during the transition in order tospin up the central rotor 140 to an appropriate rotational speed toprovide lift for forward flight. Once the aircraft 100 is established inforward flight, the rotor motor may be deactivated as the relativeairflow against the central rotor 140 becomes fast enough to causeautorotation of the central rotor 140.

The aircraft 100 may transition from forward flight to vertical flight avarious times during a mission, for example, upon arrival at a remotelocation where the aircraft 100 will hover for a period of time, uponarrival at a landing site, etc. The transition from forward flight tovertical flight begins with the aircraft 100 configured for horizontalflight. In this configuration, the proprotors 136 l, 136 r are in theforward flight position illustrated in FIG. 6 , and the central rotor140 is turning by free autorotation as the aircraft travels in forwardflight. To transition to vertical flight, the proprotors 136 l, 136 rsimultaneously tilt upward into the VTOL/hover position illustrated inFIG. 6 . The central rotor 140 is tilted forward (e.g., to a levelorientation substantially parallel to the proprotors 136 l, 136 r). Insome embodiments, the proprotors 136 l, 136 r may be tilted beyondvertical, such as between 5 degrees and 30 degrees aft of vertical, ifdesired to slow the forward airspeed of the aircraft 100. As theairspeed decreases and/or as the central rotor 140 tilts back to a levelorientation, the autorotation of the central rotor 140 may decreaseand/or stop. The rotor motor may be activated to turn the central rotor140 to provide additional lift and/or stabilization to the aircraft 100during the vertical flight phase.

Throughout the preceding disclosure, proprotors 136 l, 136 r aredescribed as being independently tiltable and controllable in order toachieve various flight control functions. FIGS. 7-10 further illustrateexample components configured to implement the proprotor controlfeatures described herein. FIG. 7 depicts a medial portion of thelateral boom 132 of an example aircraft 100 including proprotor supportand control components located thereon. FIG. 8 is a detailed view of theproprotor support and control components of FIG. 7 . As shown in FIG. 7, in some embodiments, the boom 132 may comprise a tubular medialstructure, and the distal arms 134 l, 134 r may be narrower tubularcomponents coaxially mounted partially within the tubular medialstructure. As shown in FIGS. 7 and 8 , the aircraft 100 includes agearbox 200, which may be located within the proprotor tilt controlhousing 130 illustrated in FIGS. 1-4 . As shown in FIG. 8 , the gearbox200 includes master gears 204 independently driven by proprotor tiltservos 202. The master gears 204 are positioned so as to mesh with slavegears 206, which are coaxial with and rotationally fixed to inner endsof the distal arms 134 l, 134 r. Thus, the tilt servos 202 can tilt theproprotors 136 l, 136 r by rotating the master gears 202.

As shown in FIGS. 9 and 10 , a proprotor motor 210 is disposed at anouter end of each distal arm 134 l, 134 r of the boom 132. Eachproprotor motor 210 is rotationally fixed with respect to thecorresponding distal arm 134 l, 134 r, such that the rotational axis ofthe proprotor motor 210 is perpendicular to the lateral axis along thedistal arm 134 l, 134 r. Thus, rotation of a distal arm 134 l, 134 rabout the lateral axis, under control of the proprotor tilt servos 202,produces fore and aft tilting of the rotational axis of the proprotormotor 210. Accordingly, each of the tilting operations described abovewith respect to the proprotors 136 l, 136 r may be achieved by actuatingthe proprotor tilt servos 202 individually or simultaneously.

FIGS. 11-15 illustrate an example lower rotor hub 300 of an exampleaircraft such as the aircraft 100. As described above, the lower rotorhub 300 serves as an attachment point for the upper rotor hub assembly400 and the central rotor 140. The lower rotor hub 300 is alsoconfigured to provide tilting and powered rotation of the central rotor140, as will now be described. The lower rotor hub 300 includes a rotormount shaft 302 having one or more mounting pin holes 303 extendingtherethrough, a central rotor slave gear 304 meshed with a central rotormaster gear 306, a central rotor drive motor 308, a tilt bearing 310,and a central rotor tilt servo 312.

The rotor mount shaft 302 serves as a mounting point for the centralrotor 140. The central rotor 140, not shown in FIGS. 11-15 , may bemounted to the aircraft 100 by coupling the upper rotor hub assembly 400to the rotor mount shaft 302 and securing the upper rotor hub assembly400 using one or more pins extending through mounting pin holes 303.Pin-based mounting of the upper rotor hub assembly 400 allows thecentral rotor 140 to be easily and quickly attached to or detached fromthe aircraft 100 for transportation.

The central rotor slave gear 304 is coaxial with the rotor mount shaft302 and is configured to transfer rotational motion of the central rotormaster gear 306 to the rotor mount shaft 302 to drive the central rotor140 during powered operation of the central rotor 140. In someembodiments, the central rotor slave gear 304 is coupled to the rotormount shaft 302 by a clutch mechanism and/or a one-way bearing (e.g., aone-way bearing 305) such that rotational motion of the central rotorslave gear 304 in a first direction (e.g., clockwise) is transferred tothe rotor mount shaft 302, but the rotor mount shaft 302 is free to spinin the same direction (e.g., clockwise) when the central rotor slavegear 304 is not rotating or is rotating more slowly than the rotor mountshaft 302. Thus, the central rotor drive motor 308 can power the centralrotor 140 (e.g., during vertical flight and/or during the transitionfrom vertical flight to forward flight) by turning the central rotormaster gear 306, which in turn causes the central rotor slave gear 304and rotor mount shaft 302 to rotate.

The central rotor tilt servo 312 is configured to tilt the lower rotorhub 300 relative to the mast 120. Actuation of the central rotor tiltservo 312 causes rotation of the lower rotor hub 300 about the tiltbearing 310, which accommodates motion about a lateral axisperpendicular to the rotor mount shaft 302.

FIG. 14 illustrates an example forward-most tilt position of the lowerrotor hub 300. In the tilt position illustrated in FIG. 14 , the rotormount shaft 302 is substantially aligned with a vertical axis of theaircraft 100, and a spinning central rotor 140 attached to the lowerrotor mount 300 would produce a lifting force directly upward. The tiltposition illustrated in FIG. 14 may be used, for example, duringvertical flight while the proprotors 136 l, 136 r are in the VTOL/hoverposition illustrated in FIG. 6 .

FIG. 15 illustrates an example rear-most tilt position of the lowerrotor hub 300. In the tilt position illustrated in FIG. 15 , the rotormount shaft 302 is tilted rearward approximately 20 degrees relative tovertical, such that a spinning central rotor 140 attached to the lowerrotor mount 300 would rotate within a plane tilted approximately 20degrees relative to a longitudinal axis of the aircraft 100. The tiltposition illustrated in FIG. 15 may be used, for example, duringhorizontal flight while the proprotors 136 l, 136 r are in the forwardflight position illustrated in FIG. 6 . The central rotor tilt servo 312tilts the lower rotor hub 300 between the positions illustrated in FIGS.14 and 15 when the aircraft 100 transitions from vertical flight toforward flight, or from forward flight to vertical flight, as describedabove.

FIGS. 16-20 depict an example upper rotor hub assembly such as the upperrotor hub assembly 400 of the aircraft 100. FIG. 16 depicts the upperrotor hub assembly 400 including two mounting members 402, 404coupleable to the lower rotor hub 300 of FIGS. 11-15 , as well as aninner portion of a rotor blade 142 of the central rotor 140. FIGS. 17and 18 depict the mounting members 402, 404 separately, including anexample rotor blade 142 mounted to the first mounting member 402. FIGS.19 and 20 depict the upper rotor hub assembly 400 in a foldedconfiguration.

As described above, the central rotor 140 of the aircraft 100 functionssimilarly to the rotor of a gyroplane when the aircraft 100 is in aforward flight mode. In contrast to helicopter rotors, in which bladesmay be hinged or otherwise able to move vertically and/or horizontallyrelative to the main rotor hub, gyroplane rotors perform mostefficiently when the blades are rigidly fixed relative to the centralhub and each pair of opposing blades remains symmetrically opposed.While not required, exact alignment of the blades may substantiallyimprove performance. Thus, if compact transportation of the aircraft 100is desired, it may be cumbersome to remove the blades 142 from the upperrotor hub assembly 400 for transport due to the time required tocarefully align the blades 142 when reattaching them. As will now bedescribed, the upper rotor hub assembly 400 is easily detachable fromand attachable to the lower rotor hub 300, and may be folded such thatthe entire central rotor 140 and upper rotor hub assembly 400 may betransported in a compact form while the blades 142 remain attached tothe mounting members 402, 404.

The upper rotor hub assembly 400 includes a first mounting member 402and a second mounting member 404. The mounting members 402, 404 eachinclude mounting pin holes 405 configured to align with the mounting pinholes 303 of the lower rotor hub 300 of FIGS. 11-15 . As shown in FIG.16 , the mounting members 402, 404 are shaped so as to nest togetherwhen both mounting members 402, 404 are mounted to the lower rotor hub300. In addition, the mounting members 402, 404 are shaped such that,when nested together in a mounted configuration, each mounting member402, 404 can teeter independently about the axis of its mounting pin. Insome embodiments, the aircraft 100 can be operated with only two blades142 by using only the first mounting member 402 or only the secondmounting member 404. Two-bladed operation is associated with less liftand less induced drag during forward flight, and may be desirable whenthe aircraft 100 is carrying a relatively light load and less liftingcapacity is required; four-bladed operation is associated with more liftand more induced drag during forward flight, and may be desirable whenthe aircraft 100 is carrying a relatively heavy load and greater liftingcapacity is required.

Each mounting member 402, 404 includes two mounting brackets 406disposed on opposite sides of the mounting member 402, 404 about thecenter of the hub. Each mounting bracket 406 includes two or moremounting holes 408 spaced apart to align with corresponding mountingholes 409 within the rotor blades 142. In some embodiments, more thantwo mounting holes 408, 409, such as three or more mounting holes, areprovided in order to form a structurally robust connection between therotor blades 142 and the mounting brackets 406.

The mounting members 402, 404 each include two hinges 410 disposedbetween the mounting brackets 406 and the central portion of themounting members 402, 404. The hinges 410 allow the mounting brackets406 and rotor blades 142 to be folded about the axis of the hinges 410while the aircraft is not in flight. Each mounting bracket 406 isrigidly coupled to locking plates 414 having holes located to align withlocking pin holes 413 of the central portion of the mounting members402, 404. When a hinge 410 is in the fully extended position (e.g., forflight), the hinge 410 may be locked in the fully extended position byinserting a locking pin 412 through locking pin holes 413 and theadjacent locking plates 414. The locking pins 412 may include aretaining mechanism such as spring-loaded retaining balls or the like,to prevent the locking pins 412 from pulling out during operation.

In some embodiments, instead of or in addition to being foldable, themounting members 402, 404 may also permit the rotor blades 142 to beremoved from the mounting members 402, 402 in a manner that retains thealignment of the rotor blades 142 when they are reinserted. For example,the rotor blades 142 may be slidably mounted along one or more railsdisposed within the mounting brackets 406. Each mounting bracket 406 mayinclude a release button which, when depressed, permits a rotor blade142 within the mounting bracket 406 to slide outward to be removed. Insome embodiments, the mounting brackets 406 and/or the mounting members402, 404 may include a bayonet-type mounting system which maintains theappropriate alignment between opposing rotor blades 142. Bayonet-typemounting systems will be described in greater detail with reference toFIGS. 25 and 26 .

FIGS. 19 and 20 illustrate the folded configuration of mounting members402, 404. As shown in FIGS. 19 and 20 , the removal of the locking pins412 from the mounting members 402, 404 allows the locking plates 414 tomove away from the locking pin holes 413, allowing the mounting members402, 404 to fold at the hinges 410 into a folded configuration. In thefolded configuration, both mounting brackets 406 and rotor blades 142attached to each mounting member 402, 404 are substantially parallel.Thus, a two-bladed or four-bladed central rotor 140 can be transportedwithin a relatively narrow rectangular container having a length onlyslightly longer than each of the rotor blades 142. This foldingconfiguration is substantially more efficient than transporting thecentral rotor 140 in a flight configuration, which would require acontainer having two perpendicular dimensions of at least twice thelength of each rotor blade 142. Moreover, in some cases it may bedesirable to fold the central rotor 140 while the central rotor 140remains attached to the aircraft 100. The folding mechanism describedherein permits the two or four rotor blades 142 to be folded upward suchthat the horizontal footprint of the aircraft 100 may be minimized whileit is being stored on the ground.

FIGS. 21 and 22 depict the upper rotor hub assembly 400 of FIGS. 16-20coupled to the lower rotor hub 300 of FIGS. 11-15 . Each mounting member402, 404 is coupled to the lower rotor hub 300 by a mounting pin 416passing through the mounting pin holes 405 of the mounting member 402,404 and the mounting pin holes 303 of the lower rotor hub 303. Similarto the locking pins 412 that lock the mounting members 402, 404 in thefully extended position, the mounting pins 416 may include a retainingmechanism such as spring-loaded retaining balls or the like, to preventthe mounting pins 416 from pulling out during operation.

FIGS. 23 and 24 illustrate a further example configuration of an upperrotor hub assembly 400. The example upper rotor hub assembly 400 ofFIGS. 23 and 24 further includes an upper rotor hub 416 which serves asa removable attachment point for the mounting members 402, 404. Themounting members 402, 404 may be coupled to the upper rotor hub 416 suchas by similar mounting pins 412. In some embodiments, the mounting pins412 coupling the mounting members 402, 404 to the upper rotor hub 416may be permanent or semi-permanent pins, rather than the easilyremovable pins of FIGS. 16-22 .

As shown in FIG. 24 , the upper rotor hub 416 may in turn be connectedto the rotor mount shaft 302 of the lower rotor hub by a furthermounting pin 418, which may similarly include spring-based or otherretaining elements configured to be quickly releasable. Thus, the entireupper rotor mount hub assembly 400 may be folded and subsequentlyremoved from the aircraft by removing a single mounting pin 418.

FIGS. 25 and 26 illustrate a further example configuration of an upperrotor hub assembly 400 in which a bayonet-type mounting system allowsthe rotor blades 142 to be attached and detached from the upper rotorhub assembly 400. Similar to the configuration of FIGS. 23 and 24 , theupper rotor hub assembly 400 of FIGS. 25 and 26 includes two mountingmembers 402, 404 attached to an upper rotor hub 416. Each rotor blade142 is mounted to a mounting bracket 406 by two, three, or morefasteners. Each mounting bracket 406 is fixed to a mounting body 420, aportion or all of which may be integrally formed with the mountingbracket 406. A blade attachment pin hole 422 extends laterally throughthe mounting body 420.

Each opposing end of each mounting member 402, 404 includes a mountingbody opening 424 sized and shaped to receive a mounting body 420.Additional blade attachment pin holes 426 extend laterally through thesides of the mounting members 402, 404. Thus, as illustrated by thetransition from FIG. 25 to FIG. 26 , each rotor blade 142 may be mountedby sliding the mounting body 420 into a mounting body opening 424 untilthe blade attachment pin holes 422 are substantially aligned with bladeattachment pin holes 426. A blade attachment pin 428 may then beinserted through the blade attachment pin holes 422, 426 to secure therotor blade 142 to the upper rotor hub assembly 400. Dismounting of theblades, such as for storage, transport, etc., may be accomplished byremoving the blade attachment pin 428 from the blade attachment pinholes 422, 426 and subsequently sliding the mounting body 420 out of themounting body opening 424.

As described above, it is typically desirable for the rotor to have apositive blade pitch when operating in vertical powered flight, such asin a hover or VTOL phases of flight. In contrast, when autorotation isused, such as in forward flight of a gyroplane, it is desirable for therotor blades to have a flat or zero pitch, or a substantially lesspositive pitch than in vertical flight. Accordingly, some embodiments ofthe present technology include rotor assemblies configured toselectively change the pitch angle of the rotor blades while maintainingthe teetering motion desirable for low-pitch forward flight.Advantageously, the embodiments disclosed herein accomplish blade pitchcontrol without requiring the weight and complexity of a swashplate asis typically utilized for blade pitch control in helicopters.

FIG. 27 depicts an example embodiment of a rotor hub assembly includinga teetering pivot point and synchronization linkages for controlling thepitch of the rotor blades. FIGS. 28 and 29 illustrate side andcross-sectional views of the rotor hub assembly of FIG. 27 in a zeropitch configuration. FIGS. 30 and 31 illustrate side and cross-sectionalviews of the rotor hub assembly of FIGS. 27-29 in a positive pitchconfiguration. FIG. 32 depicts an exploded view of the rotor hubassembly of FIGS. 27-31 . The rotor assembly illustrated in FIGS. 27-32may be implemented in conjunction with any of the aerial vehiclesdisclosed herein.

In order for a gyrocopter rotor assembly to generate upward thrust, thepitch of the blades is increased from zero or approximately zero to aselected pitch angle greater than zero. Additionally, the rotor isturned at an RPM that will generate enough thrust to lift the weight ofthe aircraft and payload. FIGS. 28 and 29 illustrate the assembly withthe rotor blades being at a zero-degree pitch, and FIGS. 30 and 31illustrate the assembly with the rotor blades being at a higher pitchthat creates lift.

As shown in FIGS. 28, 30, and 32 , the rotor head assembly includes ateetering pivot point attached to the rotating shaft, such that therotor head can teeter freely during forward flight. The rotor headassembly further includes synchronization linkages connecting the twoopposing arms of the rotor head that holds the two rotor blades. Theselinkages synchronize the movement of the two arms as they slide out andtwist to create the desired pitch.

Each arm of the rotor assembly includes an outer cylinder configured toretain a removable rotor blade, as described in greater detail withreference to FIGS. 25 and 26 , and an inner cylinder sized and shaped toat least partially fit within the outer cylinder. A biasing element suchas a gas spring is disposed between an end backstop of the innercylinder and an inward-facing surface of the outer cylinder arm. Theouter cylinder arm is connected on both ends to the synchronizationlinkages via arms 1 and 2 extending through arm apertures in the outersurface of the outer cylinder. Each arm passes through the two armapertures of the corresponding outer cylinder and the two rotororientation tracks of the corresponding inner cylinder, and is coupledat each end to an outer end of a synchronization linkage.

Advantageously, the rotor assembly mechanism of FIGS. 27-32 can providefor automatic adjustment of blade pitch based on the rotational speed ofthe rotor assembly, without requiring a servo or other control mechanismto adjust the blade pitch. For example, the gas spring and associatedcomponents can cause the blade pitch to increase automatically at higherRPM, and to decrease automatically at low RPM. The gas spring is ratedto collapse at a certain pound force. Thus, as the rotational speed ofthe rotor increases in a powered vertical flight mode, the correspondingincrease in centrifugal force causes the gas spring to at leastpartially collapse, letting the outer cylinder slide outwards. As theouter cylinder slides outward, the slope of the rotor orientation trackscauses the outer cylinder and attached rotor blade to twist, creating adesired pitch for the blade. The synchronization linkages simultaneouslyslide outwards and twist to the desired pitch for lift generation, andmaintain an equal or substantially equal blade pitch between the twoblades. As the RPM is subsequently reduced, the outer cylinders slideinwards through the assistance of the two gas springs, bringing theblade angle back to zero as arms 1 and 2 slide within the rotororientation tracks, allowing the gyrocopter to fly forward with the aidof the pusher motor/propeller assembly or, in any of the variousembodiments including two laterally mounted tiltable or non-tiltableproprotors (e.g., proprotors 136 l, 136 r), allowing the gyrocopter tofly forward with the aid of the left and right proprotors.

In some embodiments, the gas spring may be configured to collapse when apredetermined outward force (e.g., radially outward from the center ofthe rotor assembly toward the blade) is applied. The predetermined forcemay be selected, based at least partially on the mass of the outercylinder and blade, such that the outer cylinder collapses and increasesthe rotor blade pitch at a predetermined range of rotational speeds. Inone particular example, the predetermined force is selected such that alower RPM range, such as 250-450 RPM, does not create sufficientcentrifugal force to collapse the gas spring outward, while a higher RPMrange, such as above 550 RPM, creates sufficient centrifugal force tocause the gas spring to remain collapsed. The aircraft may thus becontrolled to operate with the main rotor turning at 450 RPM or slowerwhile in horizontal flight, and with the main rotor turning at 550 RPMor faster while in vertical or hovering flight. In some embodiments, theaircraft may be configured to avoid operating for extended periods withthe main rotor turning at speeds in an intermediate or safety RPM range(e.g., between 450 RPM and 550 RPM in the particular example above) atwhich the centrifugal force created by the rotor blades and outercylinders may be great enough to partially collapse the gas spring, butmay not be sufficient to fully collapse the gas spring.

Consistent with the automatic control of blade pitch based on rotor RPM,it may be desirable to increase and decrease the rotational speed of therotor on command. In addition, it may be desirable to increase ordecrease rotor RPM during flight in order to achieve desirable oroptimized flight characteristics. FIG. 33 illustrates an example gearingsystem for increasing the RPM of the upper rotor hub assembly of FIGS.27-32 . FIG. 34 is a cross-sectional view illustrating a braking systemfor reducing the rotational speed of the upper rotor hub assembly ofFIGS. 27-32 .

As shown in FIG. 33 , rotational speed of the rotor may be increased bya motor via a large gear having a one-way bearing, as describedelsewhere herein. When that motor is powered and turns the rotor headand blades attached to the rotor head, it generates a significant torquethat if not countered, will spin the aircraft uncontrollably. In orderto avoid undesirable spinning, the two proprotors that are located oneither side of the vehicle (e.g., as shown in FIGS. 1-4 ) at a specificCG (center of gravity) location, can operate independently at differentspeeds so as to offset the torque created by the rotor head assembly.The rotational speeds of the proprotors are variable in order to matchthe counter torque force required based on the torque created by therotor head assembly. In some embodiments, depending on the amount oftorque generated by the rotor head assembly, one proprotor (e.g., theleft proprotor) may operate at up to full speed while the otherproprotor operates at a lower RPM, at a full stop, or in a reverseddirection, in order to fully counter the torque of the rotor headassembly. Thus, the two proprotors can be independently controlled tooperate at different rotational velocities and/or in differentrotational directions. In some embodiments, the other proprotor (e.g.,the right proprotor) may even be rotated 180 degrees and operated at asuitable RPM to further provide counter torque. When the aircrafttransitions to forward flight, the proprotors can transition to turn atthe same RPM so as to provide stability and maneuverability during thehorizontal phase of flight.

Referring now to FIG. 34 , in order to slow down or control the rotorRPM, a disk brake is located at the bottom of the rotor shaft. The diskbrake is coupled with a bracket that is part of the pivoting assemblybracket and which holds the brake pads, calipers, and brake servo. Thecalipers that hold the brake pads are activated by the brake servo andare activated at any time in order to slow down the rotor or assist instopping it completely once the aircraft has landed safely. This bracketassembly allows the rotor head to move forwards and backwards asrequired for the desired flight mode.

FIGS. 35-37 illustrate an example aircraft including a tiltableempennage. In some embodiments, the empennage, including the horizontalstabilizer and one or two rudders, are turned 90 degrees downwardsduring vertical flight, thereby streamlining the airflow created by thedownwards thrust from the rotating blades of the main rotor. Theempennage may be rotatable about a lengthwise axis of the horizontalstabilizer (e.g., a lateral axis with regard to the aircraft). FIGS. 35and 36 illustrate the empennage in an upright position, such as forforward flight. FIG. 37 illustrates the empennage in a lowered position,tilted approximately 90 degrees forward or backward such that thehorizontal stabilizer is oriented substantially within a vertical plane.Once the tail assembly is in the lowered position, it has two additionalfunctions: First, by changing the angle of the stabilizer from 90degrees plus/minus, it will move the gyrocopter forwards or backwards ina controlled manner. Second, the rudders allow for additional countertorque fine tuning and being able to turn the aircraft left or right.The down turned wings have the same capability but have the primarycounter torque function. Moreover, the vertical orientation of thehorizontal stabilizer in the lowered position reduces interference withthe downward airflow created by the rotor, improving vertical and hoverflight efficiency. The rotation of the tiltable empennage betweenupright and lowered positions may be actuated by an empennage tilt servodisposed within the aircraft.

FIGS. 38-41 illustrate a further example embodiment of an aerial vehicleincluding a single proprotor and pivotable wings configured to operateas control surfaces. The aerial vehicle of FIGS. 38-41 includes acentral rotor and a single proprotor mounted to the mast so as toprovide centerline thrust for the aerial vehicle. Independentlypivotable wings are mounted at the sides of the mast. It will beunderstood that the single proprotor configuration of FIGS. 38-41 may beimplemented with any of the aerial vehicle embodiments disclosed herein.

As shown in FIGS. 38 and 39 , in horizontal forward flight, the twowings have substantially the same horizontal orientation such that thewings are streamlined for forward flight. In some embodiments, the wingsmay further be used for flight control functionality in forward flight,such as to provide a rolling or pitching moment.

As shown in FIGS. 40 and 41 , the two wings can be differentiallypositioned so as to provide a counter torque moment while the main rotoris powered in vertical or hovering flight phases. In this counter torqueposition, the left wing is pivoted by more than 90 degrees and the rightwing is pivoted by less than 90 degrees, such that the left wingdeflects the rotor downwash forward and the right wing deflects therotor downwash aft. Thus, the counter torque configuration of FIGS. 40and 41 creates a counterclockwise torque that counters the clockwisetorque created by the powered rotor.

FIGS. 42 and 43 illustrate a further example configuration of anaircraft 100 in accordance with the technology described herein. Asshown in FIGS. 42 and 43 , the aircraft 100 includes side-mountedproprotors 136 l, 136 r in a fixed (e.g., non-tiltable) configuration.Empennage 150 is fixed relative to the aircraft 100 and can be free ofcontrol surfaces such as rudders or elevators. A tilt bearing 310 isconfigured to allow fore-aft tiling, lateral tilting, and/or acombination of fore-aft and lateral tilting of the central rotor. Motorsdriving the proprotors 136 l, 136 r may be configured to individuallyand/or differentially control the speed and/or rotational direction ofproprotors 136 l, 136 r. For example, to create a yawing moment tocounter the torque created by a clockwise-rotating central rotor, theleft proprotor 136 l may be powered to rotate in a first direction at adesired rotational speed to produce forward thrust while the rightproprotor 136 r is powered to rotate in a second direction opposite thefirst direction (e.g., a reversed rotational direction) at a desiredrotational speed to produce reverse thrust, or is powered to rotate inthe first direction at a lower rotational speed relative to therotational speed of the left proprotor 136 l, such that a rightwardyawing moment is created. An opposite counter-torque scheme may beimplemented for a counterclockwise-rotating central rotor. The variablerotational speeds and/or directions of the proprotors 136 l, 136 r maybe controlled in coordination with fore-aft and/or lateral tilting ofthe central rotor at the tilt bearing 310 (e.g., under control of anautopilot or other control circuitry) to selectively control roll,pitch, yaw, and thrust of the aircraft during forward flight, asdescribed elsewhere herein.

FIGS. 44 and 45 illustrate an example rotor assembly 440 in accordancewith the present technology. An upper rotor hub assembly includes acentral section 442 and mounting brackets 444 configured as abayonet-type blade mounting system and configured to receive rotorblades therein, as described herein with reference to FIGS. 25 and 26 .A blade pitch adjustment system 446 provides for automatic adjustment ofrotor blade pitch based on rotational velocity, as described herein withreference to FIGS. 27-32 . Rotor tilt servos 448, 450, 452, and 454 areprovided to control fore-aft and/or lateral tilting of the upper rotorhub assembly. Thus, a motor 456 or corresponding lower rotor hubassembly may remain fixed relative to a fuselage of an aircraft (e.g.,aircraft 100 as described elsewhere herein) in which the rotor assembly440 is installed, while rotor tilt servos 448, 450, 452, 454 selectivelytilt the upper rotor hub assembly and connected rotor blades to controlpitch and/or roll of the aircraft. In some embodiments, the rotorassembly 440 may include only two tilt servos, for example, servos 448and 450, with one of servos 448 and 450 oriented to control fore-afttilting, and the other one of servos 448 and 450 oriented to controllateral tilting. Additional tilt servos 452, 454 may be included forredundancy and/or for additional stability when controlling tilt of theupper rotor hub assembly.

While certain embodiments have been described, these embodiments havebeen presented by way of example only, and are not intended to limit thescope of the disclosure. Indeed, the novel methods and systems describedherein may be embodied in a variety of other forms. Furthermore, variousomissions, substitutions, and changes in the systems and methodsdescribed herein may be made without departing from the spirit of thedisclosure. The accompanying claims and their equivalents are intendedto cover such forms or modifications as would fall within the scope andspirit of the disclosure. Accordingly, the scope of the presentdisclosure is defined only by reference to the appended claims.

Features, materials, characteristics, or groups described in conjunctionwith a particular aspect, embodiment, or example are to be understood tobe applicable to any other aspect, embodiment or example described inthis section or elsewhere in this specification unless incompatibletherewith. All of the features disclosed in this specification(including any accompanying claims, abstract and drawings), and/or allof the steps of any method or process so disclosed, may be combined inany combination, except combinations where at least some of suchfeatures and/or steps are mutually exclusive. The protection is notrestricted to the details of any foregoing embodiments. The protectionextends to any novel one, or any novel combination, of the featuresdisclosed in this specification (including any accompanying claims,abstract and drawings), or to any novel one, or any novel combination,of the steps of any method or process so disclosed.

Furthermore, certain features that are described in this disclosure inthe context of separate implementations can also be implemented incombination in a single implementation. Conversely, various featuresthat are described in the context of a single implementation can also beimplemented in multiple implementations separately or in any suitablesubcombination. Moreover, although features may be described above asacting in certain combinations, one or more features from a claimedcombination can, in some cases, be excised from the combination, and thecombination may be claimed as a subcombination or variation of asubcombination.

Moreover, while operations may be depicted in the drawings or describedin the specification in a particular order, such operations need not beperformed in the particular order shown or in sequential order, or thatall operations be performed, to achieve desirable results. Otheroperations that are not depicted or described can be incorporated in theexample methods and processes. For example, one or more additionaloperations can be performed before, after, simultaneously, or betweenany of the described operations. Further, the operations may berearranged or reordered in other implementations. Those skilled in theart will appreciate that in some embodiments, the actual steps taken inthe processes illustrated and/or disclosed may differ from those shownin the figures. Depending on the embodiment, certain of the stepsdescribed above may be removed, others may be added. Furthermore, thefeatures and attributes of the specific embodiments disclosed above maybe combined in different ways to form additional embodiments, all ofwhich fall within the scope of the present disclosure. Also, theseparation of various system components in the implementations describedabove should not be understood as requiring such separation in allimplementations, and it should be understood that the describedcomponents and systems can generally be integrated together in a singleproduct or packaged into multiple products.

For purposes of this disclosure, certain aspects, advantages, and novelfeatures are described herein. Not necessarily all such advantages maybe achieved in accordance with any particular embodiment. Thus, forexample, those skilled in the art will recognize that the disclosure maybe embodied or carried out in a manner that achieves one advantage or agroup of advantages as taught herein without necessarily achieving otheradvantages as may be taught or suggested herein.

Conditional language, such as “can,” “could,” “might,” or “may,” unlessspecifically stated otherwise, or otherwise understood within thecontext as used, is generally intended to convey that certainembodiments include, while other embodiments do not include, certainfeatures, elements, and/or steps. Thus, such conditional language is notgenerally intended to imply that features, elements, and/or steps are inany way required for one or more embodiments or that one or moreembodiments necessarily include logic for deciding, with or without userinput or prompting, whether these features, elements, and/or steps areincluded or are to be performed in any particular embodiment.

Conjunctive language such as the phrase “at least one of X, Y, and Z,”unless specifically stated otherwise, is otherwise understood with thecontext as used in general to convey that an item, term, etc. may beeither X, Y, or Z. Thus, such conjunctive language is not generallyintended to imply that certain embodiments require the presence of atleast one of X, at least one of Y, and at least one of Z.

The scope of the present disclosure is not intended to be limited by thespecific disclosures of preferred embodiments in this section orelsewhere in this specification, and may be defined by claims aspresented in this section or elsewhere in this specification or aspresented in the future. The language of the claims is to be interpretedbroadly based on the language employed in the claims and not limited tothe examples described in the present specification or during theprosecution of the application, which examples are to be construed asnon-exclusive.

What is claimed is:
 1. An aircraft comprising: a rotor configured toprovide lift; a first proprotor laterally displaced relative to alongitudinal axis of the aircraft and configured to produce thrustsubstantially parallel to the longitudinal axis; a second proprotorlaterally displaced relative to the longitudinal axis of the aircraft,the second proprotor disposed on an opposite side of the of thelongitudinal axis from the first proprotor and configured to producethrust substantially parallel to the longitudinal axis; one or moremotors configured to power the rotor, the first proprotor, and thesecond proprotor; and control circuitry in communication with the one ormore motors and configured to control the aircraft in a vertical takeoffor landing mode by causing the one or more motors to at least: drive therotor to provide lift; drive the first proprotor in a forward direction;and drive the second proprotor in a reversed direction, such that thefirst proprotor and the second proprotor counteract a torque created bythe rotor; wherein the control circuitry is further configured totransition the aircraft from the vertical takeoff or landing mode to aforward flight mode by at least: causing the one or more motors to stopdriving the second proprotor in the reverse direction; causing the oneor more motors to increase a rotational speed of the first proprotor inthe forward direction while a reverse thrust created by the secondproprotor decreases to zero; causing the one or more motors to drive thesecond proprotor in the forward direction to provide forward thrust; andcausing the one or more motors to stop driving the rotor such that therotor is allowed to rotate by autorotation.
 2. The aircraft of claim 1,wherein, in the vertical takeoff or landing mode, a forward thrustcreated by driving the first prop rotor in the forward direction issubstantially equal in magnitude to a reverse thrust created by drivingthe second proprotor in the reversed direction.
 3. The aircraft of claim2, wherein, in the vertical takeoff or landing mode, a rotational speedof the second proprotor in the reversed direction is greater than therotational speed of the first proprotor in the forward direction.
 4. Theaircraft of claim 1, wherein the rotor rotates in a clockwise directionwhen viewed from above, the first proprotor is disposed on a left sideof the aircraft, and the second proprotor is disposed on a right side ofthe aircraft.
 5. The aircraft of claim 1, wherein the rotor rotates in acounterclockwise direction when viewed from above, the first proprotoris disposed on a right side of the aircraft, and the second proprotor isdisposed on a left side of the aircraft.
 6. The aircraft of claim 1,wherein the control circuitry is further configured to cause the one ormore motors to decrease the rotational speed of the first proprotorafter the second proprotor begins operation in the forward direction. 7.The aircraft of claim 1, wherein the control circuitry is furtherconfigured to cause the rotor to tilt, during the transition to theforward flight mode, from a vertical position to a backward-tiltedposition that facilitates autorotation of the rotor blades.
 8. Theaircraft of claim 1, wherein the control circuitry is further configuredto control yawing of the aircraft in the forward flight mode bycontrolling rotational speeds of the first and second proprotors.
 9. Theaircraft of claim 1, wherein the control circuitry is further configuredto transition the aircraft from the forward flight mode to the verticaltakeoff or landing mode by at least: causing the one or more motors tostop driving the second proprotor in the forward direction; causing theone or more motors to drive the second proprotor in the reversedirection to provide reverse thrust; and causing the one or more motorsto drive the rotor to provide lift.
 10. The aircraft of claim 9, whereinthe control circuitry is further configured to cause the rotor to tilt,during the transition to the vertical takeoff or landing flight mode,from the backward-tilted position to the vertical position.
 11. Acomputer-implemented method for controlling an aircraft comprising arotor and at least two proprotors laterally displaced on opposing sidesof a longitudinal axis of the aircraft, the method comprising actuatingone or more motors of the aircraft to at least: drive the rotor toprovide lift; drive the first proprotor in a forward direction; anddrive the second proprotor in a reversed direction, such that the firstproprotor and the second proprotor counteract a torque created by therotor, the method further comprising transitioning the aircraft to aforward flight mode by at least: causing the one or more motors to stopdriving the second proprotor in the reverse direction; causing the oneor more motors to increase a rotational speed of the first proprotor inthe forward direction while a reverse thrust created by the secondproprotor decreases to zero; causing the one or more motors to drive thesecond proprotor in the forward direction to provide forward thrust; andcausing the one or more motors to stop driving the rotor such that therotor is allowed to rotate by autorotation.
 12. The computer-implementedmethod of claim 11, wherein a rotational speed of the first proprotorand a rotational speed of the second proprotor are controlled such thata forward thrust created by the first proprotor is substantially equalin magnitude to a reverse thrust created by the second proprotor. 13.The computer-implemented method of claim 12, wherein the rotationalspeed of the second proprotor in the reversed direction is greater thanthe rotational speed of the first proprotor in the forward direction.14. The computer-implemented method of claim 11, further comprisingcausing the one or more motors to decrease the rotational speed of thefirst proprotor after the second proprotor begins operation in theforward direction.
 15. The computer-implemented method of claim 11,further comprising, during the transition to the forward flight mode,causing the rotor to tilt from a vertical position to a backward-tiltedposition that facilitates autorotation of the rotor blades.
 16. Thecomputer-implemented method of claim 11, further comprising controllingyawing of the aircraft in the forward flight mode by controllingrotational speeds of the first and second proprotors.